US8118547B1 - Turbine inter-stage gap cooling arrangement - Google Patents

Turbine inter-stage gap cooling arrangement Download PDF

Info

Publication number
US8118547B1
US8118547B1 US12/423,874 US42387409A US8118547B1 US 8118547 B1 US8118547 B1 US 8118547B1 US 42387409 A US42387409 A US 42387409A US 8118547 B1 US8118547 B1 US 8118547B1
Authority
US
United States
Prior art keywords
blade outer
cooling air
seal
outer air
air seal
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Fee Related, expires
Application number
US12/423,874
Inventor
George Liang
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Florida Turbine Technologies Inc
Original Assignee
Florida Turbine Technologies Inc
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Florida Turbine Technologies Inc filed Critical Florida Turbine Technologies Inc
Priority to US12/423,874 priority Critical patent/US8118547B1/en
Application granted granted Critical
Publication of US8118547B1 publication Critical patent/US8118547B1/en
Assigned to FLORIDA TURBINE TECHNOLOGIES, INC. reassignment FLORIDA TURBINE TECHNOLOGIES, INC. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: LIANG, GEORGE
Assigned to SUNTRUST BANK reassignment SUNTRUST BANK SUPPLEMENT NO. 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT Assignors: CONSOLIDATED TURBINE SPECIALISTS LLC, ELWOOD INVESTMENTS LLC, FLORIDA TURBINE TECHNOLOGIES INC., FTT AMERICA, LLC, KTT CORE, INC., S&J DESIGN LLC, TURBINE EXPORT, INC.
Assigned to FLORIDA TURBINE TECHNOLOGIES, INC., CONSOLIDATED TURBINE SPECIALISTS, LLC, KTT CORE, INC., FTT AMERICA, LLC reassignment FLORIDA TURBINE TECHNOLOGIES, INC. RELEASE BY SECURED PARTY (SEE DOCUMENT FOR DETAILS). Assignors: TRUIST BANK (AS SUCCESSOR BY MERGER TO SUNTRUST BANK), COLLATERAL AGENT
Expired - Fee Related legal-status Critical Current
Adjusted expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the present invention relates generally to a gas turbine engine, and more specifically to a turbine interstage gap between a blade outer air seal and an endwall of an adjacent stator vane.
  • a gas turbine engine such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.
  • IGT industrial gas turbine
  • the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine.
  • the turbine inlet temperature is limited to the material properties of the turbine parts, such as the first stage guide vanes and rotor blades.
  • the turbine inlet temperature is limited to an amount of cooling that can be produced on a turbine vane or blade.
  • Improved cooling capability will also allow for the turbine airfoils to be exposed to higher temperatures. Improved cooling will also allow for longer part life which results in longer engine run times or longer periods between engine breakdowns.
  • FIG. 1 shows a prior art turbine vane with a bow wave effect located upstream of the turbine vanes.
  • the high pressure upstream of the vane leading edge is greater than the pressure inside the cavity formed by the gap.
  • the hot gas will flow radially inward into the cavity.
  • the ingested hot gas flows through the gap circumferentially inside the cavity towards the lower pressure zones.
  • the hot gas then flows out at locations where the cavity pressure is higher than the local hot gas pressure.
  • FIG. 2 shows a prior art turbine with a first stage rotor blade located upstream from a row of second stage stator vanes.
  • An interstage gap is formed between a blade ring for the rotor blade and a blade ring for the stator vane.
  • This arrangement in FIG. 2 includes a rotor blade 27 with a tip that forms a seal with a blade outer air seal (or BOAS) 24 , the BOAS 24 is supported by hooks on an isolation ring 22 on a forward side and a blade ring 21 on an isolation ring 25 on the aft side.
  • a first blade ring 21 supports both isolation rings 22 and 25 and includes a cooling air passage that delivers cooling air to an impingement plate 23 that includes impingement holes 28 that discharge jets of impingement cooling air onto a top surface of the BOAS.
  • An adjacent stator vane assembly includes a second blade ring 26 that supports a guide vane 11 with an outer endwall 12 .
  • an interstage gap 29 is formed between the isolation ring 25 and the vane outer diameter endwall 12 in which the hot gas ingress can occur due to the pressure differential described above.
  • the size of the bow wave is a strong function of the vane leading edge diameter and distance of the vane leading edge to the endwall edge.
  • the pressure variation in the tangential direction with the gap is sinusoidal.
  • the amount of hot gas flow penetrating the axial gap increases linearly with the increasing axial gap width. It is therefore necessary to reduce the axial gap width to a minimum allowable by tolerance limits in order to reduce the hot gas ingress.
  • a row of cooling air holes are located on the BOAS upstream from the vane leading edge diameter that discharges cooling air into the airfoil leading edge section.
  • FIG. 1 shows a cross section side view of a prior art turbine stator vane with the hot gas flow pattern and hot gas ingress flow into the outer diameter endwall and inner diameter endwall of the vane.
  • FIG. 2 shows a cross section side view of an inter-stage seal arrangement for a prior art turbine rotor blade and adjacent stator vane design with an interstage gap.
  • FIG. 3 shows a cross section side view of an inter-stage seal arrangement of the present invention for the turbine rotor blade and adjacent stator vane with an inter-stage gap.
  • FIG. 4 shows a detailed close-up view of the BOAS cooling air holes for the gap of FIG. 3 .
  • the present invention is a turbine interstage gap cooling apparatus and method for an industrial gas turbine engine that can also be used in an aero engine for the same purpose.
  • FIG. 3 shows a stage of rotor blades adjacent to an upstream from a stage of guide vanes.
  • the rotor blade 27 includes a tip that forms a seal with the BOAS 24 as in the prior art FIG. 2 .
  • the same parts in FIG. 3 are labeled as the same reference numbers as in the prior art FIG. 2 arrangement.
  • the blade outer air seal (BOAS) 24 in the FIG. 3 invention includes a row of cooling air holes 31 as seen in FIG.
  • the BOAS 24 includes an outward extending ledge 36 on the aft side that extends beyond the plane of the aft side that is flush with the isolation ring 25 as is the case in the prior art FIG. 2 BOAS.
  • the cooling air holes 31 are located above the ledge 36 and are directed to discharge the cooling air toward the transition between the concave shaped outer diameter endwall 12 and the leading edge of the airfoil 11 .
  • the cooling air holes 31 extend along the aft side of the BOAS.
  • a TBC is shown applied to the inner surface of the BOAS.
  • a tangent line 32 is tangent to the concave shaped endwall surface as seen in FIG. 4 .
  • An arrow 33 represents the direction of the hot gas flow through the vane.
  • the angle of the cooling air holes 31 and therefore the angle of injection of the cooling air 34 is half the difference between the two angles of the tangent 32 and the hot gas flow 33 .
  • the injection of the spent cooling air from the blade outer air seal trailing edge cooling through the row of metering holes 31 and into the vane leading edge nose region will eliminate the hot gas ingestion into the gap 29 that is present in the prior art inter-stage seal gap design.
  • the spent cooling air form the blade outer air seal is discharged into the vane leading edge in-between the angle formed by the streamline of the hot gas flow and a tangent to the endwall corner diameter of the vane.
  • This precise position of the spent cooling air discharge cooling holes 31 will provide proper cooling for the vane bow wave region in addition to prevent ingress of the hot gas into the gap 29 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A turbine inter-stage gap cooling and sealing arrangement for a turbine in which the blade outer air seal that forms a seal with a stage of rotor blades includes a row of cooling air holes on the back side of the blade outer air seal to discharge cooling air toward a transition between a vane endwall and the vane airfoil such that hot gas flow is not ingested into the gap formed between the BOAS and the vane endwall. The cooling air holes in the BOAS are connected to the impingement cavity on the outer surface of the BOAS to use spent impingement cooling air for discharging toward the inter-stage gap. The BOAS also includes an aft extending ledge that extends toward the vane airfoil in which the cooling air holes are located above.

Description

GOVERNMENT LICENSE RIGHTS
None.
CROSS-REFERENCE TO RELATED APPLICATIONS
None.
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine interstage gap between a blade outer air seal and an endwall of an adjacent stator vane.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.
It is well known that the efficiency of the engine can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine parts, such as the first stage guide vanes and rotor blades. Also, the turbine inlet temperature is limited to an amount of cooling that can be produced on a turbine vane or blade. Improved cooling capability will also allow for the turbine airfoils to be exposed to higher temperatures. Improved cooling will also allow for longer part life which results in longer engine run times or longer periods between engine breakdowns.
Another problem with the turbines is hot flow ingestion into a section of the turbine that is sensitive to the high temperatures such as the rim cavities or interstage gaps. Bow wave driven hot gas flow ingestion is created when the hot gas core flow enters a vane row where a leading edge of the vane induces a local blockage and thus creates a circumferential pressure variation at an intersection of the airfoil leading edge location of the vane. The leading edge of a turbine vane generates upstream pressure variations which can lead to hot gas ingress into the front gap. If proper cooling or design measures are not undertaken to prevent this hot gas ingress, exposure to the hot gas can result in severe damage to the front edges of the vane endwall as well as the turbine components located upstream of the endwall. FIG. 1 shows a prior art turbine vane with a bow wave effect located upstream of the turbine vanes. The high pressure upstream of the vane leading edge is greater than the pressure inside the cavity formed by the gap. As a result of the pressure differential, the hot gas will flow radially inward into the cavity. The ingested hot gas flows through the gap circumferentially inside the cavity towards the lower pressure zones. The hot gas then flows out at locations where the cavity pressure is higher than the local hot gas pressure.
FIG. 2 shows a prior art turbine with a first stage rotor blade located upstream from a row of second stage stator vanes. An interstage gap is formed between a blade ring for the rotor blade and a blade ring for the stator vane. This arrangement in FIG. 2 includes a rotor blade 27 with a tip that forms a seal with a blade outer air seal (or BOAS) 24, the BOAS 24 is supported by hooks on an isolation ring 22 on a forward side and a blade ring 21 on an isolation ring 25 on the aft side. A first blade ring 21 supports both isolation rings 22 and 25 and includes a cooling air passage that delivers cooling air to an impingement plate 23 that includes impingement holes 28 that discharge jets of impingement cooling air onto a top surface of the BOAS.
An adjacent stator vane assembly includes a second blade ring 26 that supports a guide vane 11 with an outer endwall 12. an interstage gap 29 is formed between the isolation ring 25 and the vane outer diameter endwall 12 in which the hot gas ingress can occur due to the pressure differential described above.
In general, the size of the bow wave is a strong function of the vane leading edge diameter and distance of the vane leading edge to the endwall edge. The pressure variation in the tangential direction with the gap is sinusoidal. The amount of hot gas flow penetrating the axial gap increases linearly with the increasing axial gap width. It is therefore necessary to reduce the axial gap width to a minimum allowable by tolerance limits in order to reduce the hot gas ingress.
As a result of the design of FIG. 2, hot gas flows in and out along the inter-stage gaps and an over-temperature occurs at the blade outer air seal edges and the blade isolation ring corresponding to the hot gas injection location. This over-temperature issue is more pronounced when an insufficient amount of inter-stage gap purge air for the axial gap is available when a strong bow wave is induced by the low solidity vane airfoil creates a high circumferential pressure variation which acts to push the mainstream hot gas into the inter-stage gap 29.
BRIEF SUMMARY OF THE INVENTION
It is an object of the present invention to provide for a turbine with an interstage gap in which the hot gas ingress into the gap is eliminated.
It is another object of the present invention to eliminate the ingress of hot gas flow caused by a differential pressure between the hot gas pressure and the cavity pressure from the bow-wave effect.
These objectives and more can be achieved by the turbine inter-stage gap cooling apparatus and method of the present invention. A row of cooling air holes are located on the BOAS upstream from the vane leading edge diameter that discharges cooling air into the airfoil leading edge section. The forced injection of the cooling air flow with the use of the blade outer air seal spent cooling air into the transition space between the vane leading edge airfoil and the vane outer diameter endwall will prevent the hot gas flow from ingesting into the interstage gap.
BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS
FIG. 1 shows a cross section side view of a prior art turbine stator vane with the hot gas flow pattern and hot gas ingress flow into the outer diameter endwall and inner diameter endwall of the vane.
FIG. 2 shows a cross section side view of an inter-stage seal arrangement for a prior art turbine rotor blade and adjacent stator vane design with an interstage gap.
FIG. 3 shows a cross section side view of an inter-stage seal arrangement of the present invention for the turbine rotor blade and adjacent stator vane with an inter-stage gap.
FIG. 4 shows a detailed close-up view of the BOAS cooling air holes for the gap of FIG. 3.
 DETAILED DESCRIPTION OF THE INVENTION
The present invention is a turbine interstage gap cooling apparatus and method for an industrial gas turbine engine that can also be used in an aero engine for the same purpose. FIG. 3 shows a stage of rotor blades adjacent to an upstream from a stage of guide vanes. The rotor blade 27 includes a tip that forms a seal with the BOAS 24 as in the prior art FIG. 2. The same parts in FIG. 3 are labeled as the same reference numbers as in the prior art FIG. 2 arrangement. The blade outer air seal (BOAS) 24 in the FIG. 3 invention includes a row of cooling air holes 31 as seen in FIG. 4 that connect the inner side of the BOAS to the aft side of the BOAS 24 such that spent impingement cooling air from the inner surface of the BOAS 24 will be discharged in the direction of the arrow shown in FIG. 4. The BOAS 24 includes an outward extending ledge 36 on the aft side that extends beyond the plane of the aft side that is flush with the isolation ring 25 as is the case in the prior art FIG. 2 BOAS. The cooling air holes 31 are located above the ledge 36 and are directed to discharge the cooling air toward the transition between the concave shaped outer diameter endwall 12 and the leading edge of the airfoil 11. The cooling air holes 31 extend along the aft side of the BOAS. A TBC is shown applied to the inner surface of the BOAS. A tangent line 32 is tangent to the concave shaped endwall surface as seen in FIG. 4. An arrow 33 represents the direction of the hot gas flow through the vane. The angle of the cooling air holes 31 and therefore the angle of injection of the cooling air 34 is half the difference between the two angles of the tangent 32 and the hot gas flow 33.
The injection of the spent cooling air from the blade outer air seal trailing edge cooling through the row of metering holes 31 and into the vane leading edge nose region will eliminate the hot gas ingestion into the gap 29 that is present in the prior art inter-stage seal gap design. The spent cooling air form the blade outer air seal is discharged into the vane leading edge in-between the angle formed by the streamline of the hot gas flow and a tangent to the endwall corner diameter of the vane. This precise position of the spent cooling air discharge cooling holes 31 will provide proper cooling for the vane bow wave region in addition to prevent ingress of the hot gas into the gap 29.

Claims (11)

I claim the following:
1. A gas turbine engine comprising:
a blade outer air seal that forms a seal with a stage or rotor blades;
a stator vane located adjacent to and downstream from the stage of rotor blades;
the stator vane having a vane airfoil extending from an outer diameter endwall;
a turbine inter-stage gap formed between the blade outer air seal and the vane outer diameter endwall in which a hot gas flow from the turbine can be ingested into; and,
a row of cooling air holes in the blade outer air seal directed to discharge cooling air at a location upstream from the inter-stage gap to prevent ingestion of the hot gas flow from the turbine.
2. The gas turbine engine of claim 1, and further comprising:
the vane endwall has a concave curvature that forms a tangent line;
the hot gas flow passes through the turbine in a specific direction; and,
the cooling holes in the blade outer air seal are angled at around one half a difference between the tangent line and the hot gas flow specific direction.
3. The gas turbine engine of claim 1, and further comprising:
the blade outer air seal includes a ledge on the aft side that extends toward the vane airfoil; and,
the cooling air holes discharge the cooling air above the ledge.
4. The gas turbine engine of claim 1, and further comprising:
the cooling air holes extend along from one side of the back side to the opposite side of the back side of the blade outer air seal.
5. The gas turbine engine of claim 1, and further comprising:
the cooling air holes open into the inner surface of the blade outer air seal such that spent impingement cooling air for the blade outer air seal flows through the cooling air holes.
6. A blade outer air seal used for form a seal between a turbine rotor blade in a gas turbine engine, the blade outer air seal comprising:
an inner surface that forms a gap with a blade tip of a turbine rotor blade;
a forward hook that secures a forward side of the blade outer air seal to a first isolation ring;
an aft hook that secures an aft side of the blade outer air seal to a second isolation ring;
an impingement cavity formed on the outer side of the blade outer air seal; and,
a row of cooling air holes that open onto a backside of the blade outer air seal and air connected to the impingement cavity.
7. The blade outer air seal of claim 6, and further comprising:
a ledge extending out from a backside of the blade outer air seal and being flush with the inner surface; and,
the row of cooling air holes opening above the ledge.
8. The blade outer air seal of claim 6, and further comprising:
the row of cooling air holes discharging cooling air at an angle slightly downward in a direction of a rotational axis of the rotor blades.
9. The blade outer air seal of claim 6, and further comprising:
the row of cooling air holes is angled to discharge jets of cooling air toward a transition between a vane endwall and an airfoil extending from the vane endwall.
10. A process for reducing an ingestion of a hot gas flow into an interstage gap formed between a stage of rotor blades and an adjacent stage of stator vanes within a gas turbine engine, the process comprising the steps of:
Impinging cooling air onto a backside surface of a blade outer air seal that forms a seal with the stage of rotor blades; and,
Discharging spent impingement cooling air from the blade outer air seal toward an upstream end of the interstage gap to prevent a hot gas flow from ingesting into the gap.
11. The process for reducing an ingestion of a hot gas flow into an interstage gap of claim 10, and further comprising the step of:
Forming a ledge on the aft side of the blade outer air seal that extends toward the vane airfoil and is located below the discharge of the spent cooling air.
US12/423,874 2009-04-15 2009-04-15 Turbine inter-stage gap cooling arrangement Expired - Fee Related US8118547B1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US12/423,874 US8118547B1 (en) 2009-04-15 2009-04-15 Turbine inter-stage gap cooling arrangement

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/423,874 US8118547B1 (en) 2009-04-15 2009-04-15 Turbine inter-stage gap cooling arrangement

Publications (1)

Publication Number Publication Date
US8118547B1 true US8118547B1 (en) 2012-02-21

Family

ID=45571953

Family Applications (1)

Application Number Title Priority Date Filing Date
US12/423,874 Expired - Fee Related US8118547B1 (en) 2009-04-15 2009-04-15 Turbine inter-stage gap cooling arrangement

Country Status (1)

Country Link
US (1) US8118547B1 (en)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8998572B2 (en) 2012-06-04 2015-04-07 United Technologies Corporation Blade outer air seal for a gas turbine engine
US9115596B2 (en) 2012-08-07 2015-08-25 United Technologies Corporation Blade outer air seal having anti-rotation feature
US20160160760A1 (en) * 2013-03-15 2016-06-09 United Technologies Corporation Self-opening cooling passages for a gas turbine engine
US9506367B2 (en) 2012-07-20 2016-11-29 United Technologies Corporation Blade outer air seal having inward pointing extension
US9617866B2 (en) 2012-07-27 2017-04-11 United Technologies Corporation Blade outer air seal for a gas turbine engine
US9689276B2 (en) 2014-07-18 2017-06-27 Pratt & Whitney Canada Corp. Annular ring assembly for shroud cooling
US9797262B2 (en) 2013-07-26 2017-10-24 United Technologies Corporation Split damped outer shroud for gas turbine engine stator arrays
US9803491B2 (en) 2012-12-31 2017-10-31 United Technologies Corporation Blade outer air seal having shiplap structure
US9845696B2 (en) 2014-12-15 2017-12-19 Pratt & Whitney Canada Corp. Turbine shroud sealing architecture
US20190063323A1 (en) * 2017-08-30 2019-02-28 United Technologies Corporation Conformal seal bow wave cooling
US10316683B2 (en) 2014-04-16 2019-06-11 United Technologies Corporation Gas turbine engine blade outer air seal thermal control system
US11156115B2 (en) 2018-06-28 2021-10-26 MTU Aero Engines AG Guide vane arrangement for turbomachine

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5752802A (en) * 1996-12-19 1998-05-19 Solar Turbines Incorporated Sealing apparatus for airfoils of gas turbine engines
US6077035A (en) * 1998-03-27 2000-06-20 Pratt & Whitney Canada Corp. Deflector for controlling entry of cooling air leakage into the gaspath of a gas turbine engine
US20050111966A1 (en) * 2003-11-26 2005-05-26 Metheny Alfred P. Construction of static structures for gas turbine engines
US20070059158A1 (en) * 2005-09-12 2007-03-15 United Technologies Corporation Turbine cooling air sealing
US20080112793A1 (en) * 2006-11-10 2008-05-15 General Electric Company Interstage cooled turbine engine
US20100254806A1 (en) * 2009-04-06 2010-10-07 General Electric Company Methods, systems and/or apparatus relating to seals for turbine engines
US20110129342A1 (en) * 2009-11-30 2011-06-02 Honeywell International Inc. Turbine assemblies with impingement cooling

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5752802A (en) * 1996-12-19 1998-05-19 Solar Turbines Incorporated Sealing apparatus for airfoils of gas turbine engines
US6077035A (en) * 1998-03-27 2000-06-20 Pratt & Whitney Canada Corp. Deflector for controlling entry of cooling air leakage into the gaspath of a gas turbine engine
US20050111966A1 (en) * 2003-11-26 2005-05-26 Metheny Alfred P. Construction of static structures for gas turbine engines
US20070059158A1 (en) * 2005-09-12 2007-03-15 United Technologies Corporation Turbine cooling air sealing
US20080112793A1 (en) * 2006-11-10 2008-05-15 General Electric Company Interstage cooled turbine engine
US7870742B2 (en) * 2006-11-10 2011-01-18 General Electric Company Interstage cooled turbine engine
US20100254806A1 (en) * 2009-04-06 2010-10-07 General Electric Company Methods, systems and/or apparatus relating to seals for turbine engines
US20110129342A1 (en) * 2009-11-30 2011-06-02 Honeywell International Inc. Turbine assemblies with impingement cooling

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8998572B2 (en) 2012-06-04 2015-04-07 United Technologies Corporation Blade outer air seal for a gas turbine engine
US9506367B2 (en) 2012-07-20 2016-11-29 United Technologies Corporation Blade outer air seal having inward pointing extension
US10436054B2 (en) 2012-07-27 2019-10-08 United Technologies Corporation Blade outer air seal for a gas turbine engine
US9617866B2 (en) 2012-07-27 2017-04-11 United Technologies Corporation Blade outer air seal for a gas turbine engine
US9115596B2 (en) 2012-08-07 2015-08-25 United Technologies Corporation Blade outer air seal having anti-rotation feature
US9803491B2 (en) 2012-12-31 2017-10-31 United Technologies Corporation Blade outer air seal having shiplap structure
US10006367B2 (en) * 2013-03-15 2018-06-26 United Technologies Corporation Self-opening cooling passages for a gas turbine engine
US20160160760A1 (en) * 2013-03-15 2016-06-09 United Technologies Corporation Self-opening cooling passages for a gas turbine engine
US9797262B2 (en) 2013-07-26 2017-10-24 United Technologies Corporation Split damped outer shroud for gas turbine engine stator arrays
US10316683B2 (en) 2014-04-16 2019-06-11 United Technologies Corporation Gas turbine engine blade outer air seal thermal control system
US9689276B2 (en) 2014-07-18 2017-06-27 Pratt & Whitney Canada Corp. Annular ring assembly for shroud cooling
US10746048B2 (en) 2014-07-18 2020-08-18 Pratt & Whitney Canada Corp. Annular ring assembly for shroud cooling
US9845696B2 (en) 2014-12-15 2017-12-19 Pratt & Whitney Canada Corp. Turbine shroud sealing architecture
US10533444B2 (en) 2014-12-15 2020-01-14 Pratt & Whitney Canada Corp. Turbine shroud sealing architecture
US20190063323A1 (en) * 2017-08-30 2019-02-28 United Technologies Corporation Conformal seal bow wave cooling
US10738701B2 (en) * 2017-08-30 2020-08-11 Raytheon Technologies Corporation Conformal seal bow wave cooling
US11156115B2 (en) 2018-06-28 2021-10-26 MTU Aero Engines AG Guide vane arrangement for turbomachine

Similar Documents

Publication Publication Date Title
US8118547B1 (en) Turbine inter-stage gap cooling arrangement
US8240980B1 (en) Turbine inter-stage gap cooling and sealing arrangement
JP4876043B2 (en) Flared tip turbine blade
US8979481B2 (en) Turbine bucket angel wing features for forward cavity flow control and related method
US8221055B1 (en) Turbine stator vane with endwall cooling
CN103075198B (en) Turbine bucket platform leading edge and associated method
US9631557B2 (en) Sealing device for a turbomachine turbine nozzle
US8834122B2 (en) Turbine bucket angel wing features for forward cavity flow control and related method
CN102187061B (en) Gas turbine
EP2644836B1 (en) Gas turbine assembly having an effusion cooled shroud segment with an abradable coating
US8967973B2 (en) Turbine bucket platform shaping for gas temperature control and related method
JP6888907B2 (en) gas turbine
US8573925B2 (en) Cooled component for a gas turbine engine
JP2007120501A (en) Interstage seal, turbine blade, and interface seal between cooled rotor and stator of gas turbine engine
US8382424B1 (en) Turbine vane mate face seal pin with impingement cooling
US9976423B2 (en) Airfoil showerhead pattern apparatus and system
US8118554B1 (en) Turbine vane with endwall cooling
US20170175557A1 (en) Gas turbine sealing
US20130318982A1 (en) Turbine cooling apparatus
JP2019056366A (en) Shield for turbine engine airfoil
JP5770970B2 (en) Turbine nozzle for gas turbine engine
EP3543465B1 (en) Blade having a tip cooling cavity and method of making same
US10301967B2 (en) Incident tolerant turbine vane gap flow discouragement
US8459933B1 (en) Turbine vane with endwall cooling
WO2019040316A1 (en) Turbine blade with leading edge showerhead hole arrangement

Legal Events

Date Code Title Description
STCF Information on status: patent grant

Free format text: PATENTED CASE

AS Assignment

Owner name: FLORIDA TURBINE TECHNOLOGIES, INC., FLORIDA

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:LIANG, GEORGE;REEL/FRAME:028242/0013

Effective date: 20120210

FPAY Fee payment

Year of fee payment: 4

SULP Surcharge for late payment
AS Assignment

Owner name: SUNTRUST BANK, GEORGIA

Free format text: SUPPLEMENT NO. 1 TO AMENDED AND RESTATED INTELLECTUAL PROPERTY SECURITY AGREEMENT;ASSIGNORS:KTT CORE, INC.;FTT AMERICA, LLC;TURBINE EXPORT, INC.;AND OTHERS;REEL/FRAME:048521/0081

Effective date: 20190301

FEPP Fee payment procedure

Free format text: MAINTENANCE FEE REMINDER MAILED (ORIGINAL EVENT CODE: REM.); ENTITY STATUS OF PATENT OWNER: SMALL ENTITY

LAPS Lapse for failure to pay maintenance fees

Free format text: PATENT EXPIRED FOR FAILURE TO PAY MAINTENANCE FEES (ORIGINAL EVENT CODE: EXP.); ENTITY STATUS OF PATENT OWNER: SMALL ENTITY

STCH Information on status: patent discontinuation

Free format text: PATENT EXPIRED DUE TO NONPAYMENT OF MAINTENANCE FEES UNDER 37 CFR 1.362

FP Lapsed due to failure to pay maintenance fee

Effective date: 20200221

AS Assignment

Owner name: FLORIDA TURBINE TECHNOLOGIES, INC., FLORIDA

Free format text: RELEASE BY SECURED PARTY;ASSIGNOR:TRUIST BANK (AS SUCCESSOR BY MERGER TO SUNTRUST BANK), COLLATERAL AGENT;REEL/FRAME:059619/0336

Effective date: 20220330

Owner name: CONSOLIDATED TURBINE SPECIALISTS, LLC, OKLAHOMA

Free format text: RELEASE BY SECURED PARTY;ASSIGNOR:TRUIST BANK (AS SUCCESSOR BY MERGER TO SUNTRUST BANK), COLLATERAL AGENT;REEL/FRAME:059619/0336

Effective date: 20220330

Owner name: FTT AMERICA, LLC, FLORIDA

Free format text: RELEASE BY SECURED PARTY;ASSIGNOR:TRUIST BANK (AS SUCCESSOR BY MERGER TO SUNTRUST BANK), COLLATERAL AGENT;REEL/FRAME:059619/0336

Effective date: 20220330

Owner name: KTT CORE, INC., FLORIDA

Free format text: RELEASE BY SECURED PARTY;ASSIGNOR:TRUIST BANK (AS SUCCESSOR BY MERGER TO SUNTRUST BANK), COLLATERAL AGENT;REEL/FRAME:059619/0336

Effective date: 20220330